Translating inlet for adjusting airflow distortion in gas turbine engine

ABSTRACT

Systems and methods for adjusting airflow distortion in a gas turbine engine using a translating inlet assembly are provided. In one embodiment, a core engine of a gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include an inner flowpath surface. A core casing can enclose the core engine. A forward end of the core casing can include a translating inlet assembly moveable between a first position and a second position. The translating inlet assembly and the inner flowpath surface can together define an inlet to an engine airflow path. A translating inlet assembly can define a first inlet area in the first position and a second inlet area in the second position, the first inlet area being greater than the second inlet area.

FIELD OF THE INVENTION

The present subject matter relates generally to gas turbine engines andmore particularly to a translating inlet for adjusting airflowdistortion in a gas turbine engine.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a core having, in serial floworder, a compressor section, a combustion section, a turbine section,and an exhaust section. During operation, an engine airflow is providedto an inlet of the compressor section where one or more axialcompressors progressively compress the air until it reaches thecombustion section. Fuel is mixed with the compressed air and burnedwithin the combustion section to provide combustion gases. Thecombustion gases are routed from the combustion section to the turbinesection. The flow of combustion gasses through the turbine sectiondrives the compressor section and is then routed through the exhaustsection, e.g., to atmosphere.

During operation, the gas turbine engine may encounter airflowdistortion in the engine airflow path upstream of the compressorsection, such as a circumferential or local flow disruption due to theangle of attack of the gas turbine engine, a cross wind, or any otherinlet anomaly. Airflow distortion can be so uneven during operation asto put portions of the compressor section at or below proper stallpressure ratios. In many cases, sufficient stall margin should bemaintained in the compressor section in order to prevent stallconditions from occurring during operation of the gas turbine engine.

One approach to maintaining a desired stall margin in a gas turbineengine is to close the variable guide vanes at the inlet to thecompressor section, thereby reducing air flow and pressure in thecompressor section below a pressure sufficient to cause stallconditions. However, closing the variable guide vanes can decrease theoverall efficiency of the gas turbine engine.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

One example aspect of the present disclosure is directed to a coreengine of a gas turbine engine. The core engine can include a compressorsection, a combustion section, and a turbine section in series flow anddefining at least in part an engine airflow path for the gas turbineengine. The core engine can also include an inner flowpath surfacepositioned at least partially within the compressor section and definingat least in part the engine airflow path. The core engine can alsoinclude a core casing at least partially enclosing the compressorsection and defining a forward end. The core casing can include atranslating inlet assembly at the forward end. The translating inletassembly and inner flowpath surface can together define an inlet to thecompressor section. The translating inlet assembly can be moveablebetween a first position defining a first inlet area at the inlet and asecond position defining a second inlet area at the inlet, with thefirst inlet area being greater than the second inlet area.

Another example aspect of the present disclosure is directed to a methodfor adjusting airflow distortion in a gas turbine engine. The gasturbine engine can include a compressor section, a combustion section,and a turbine section in series flow. The compressor section, combustionsection, and turbine section can define at least in part an engineairflow path. The gas turbine engine can include an inner flow pathsurface positioned at least partially within the compressor section anddefining at least in part the engine airflow path. The gas turbineengine can include a core casing at least partially enclosing thecompressor section and defining a forward end. The method includesdetermining, by one or more control devices, an airflow distortioncondition associated with the engine airflow path. The method can alsoinclude controlling, by the one or more control devices, a translatinginlet assembly to adjust the airflow distortion condition of the gasturbine engine. The core casing can include the translating inletassembly at the forward end. The translating inlet assembly and innerflow path surface can together define an inlet to the compressorsection. The translating inlet assembly can be moveable between a firstposition defining a first inlet area and a second position defining asecond inlet area, with the first inlet area being greater than thesecond inlet area.

Other example aspects of the present disclosure are directed to gasturbine engines, devices, apparatus, and other systems configured toadjust airflow distortion in the airflow path of a gas turbine engine.Variations and modifications can be made to these example aspects of thepresent disclosure.

These and other features, aspects and advantages of various embodimentswill become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the present disclosure and, together with thedescription, serve to explain the related principles.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic, cross-sectional view of a gas turbine engineaccording to example embodiments of the present subject matter.

FIG. 2 is schematic, cross-sectional view of a forward end of an examplegas turbine engine according to example embodiments of the presentdisclosure.

FIG. 3 is schematic, cross-sectional view of a forward end of an examplegas turbine engine according to example embodiments of the presentdisclosure.

FIG. 4 is a schematic, axial view of a forward end of an example gasturbine engine according to example embodiments of the presentdisclosure.

FIG. 5 is a schematic, axial view of a forward end of an example gasturbine engine according to example embodiments of the presentdisclosure.

FIG. 6 is schematic, cross-sectional view of a forward end of an examplegas turbine engine according to example embodiments of the presentdisclosure.

FIG. 7 is schematic, cross-sectional view of a forward end of an examplegas turbine engine according to example embodiments of the presentdisclosure.

FIG. 8 is a schematic, axial view of an array of instrumented guidevanes in an example gas turbine engine according to example embodimentsof the present disclosure.

FIG. 9 is a schematic of an individual instrumented guide vane in anexample gas turbine engine according to example embodiments of thepresent disclosure.

FIG. 10 depicts an example control device used in a control systemaccording to example embodiments of the present disclosure.

FIG. 11 depicts a flow diagram of an example method according to exampleembodiments of the present disclosure.

DETAILED DESCRIPTION OF THE INVENTION

Reference now will be made in detail to embodiments of the presentdisclosure, one or more examples of which are illustrated in thedrawings. Each example is provided by way of explanation of theinvention, not limitation of the invention. In fact, it will be apparentto those skilled in the art that various modifications and variationscan be made in the present invention without departing from the scope orspirit of the present disclosure. For instance, features illustrated ordescribed as part of one embodiment can be used with another embodimentto yield a still further embodiment. Thus, it is intended that thepresent invention covers such modifications and variations as comewithin the scope of the appended claims and their equivalents.

Example aspects of the present disclosure are directed to adjustingairflow distortion in a gas turbine engine in real-time. Airflowdistortion can be so uneven during operation of the gas turbine engineas to put portions of the compressor section at or below proper stallpressure ratios, thereby reducing the operability of the gas turbineengine. Increased stall margin headroom can therefore be necessary toaccount for airflow distortion.

The gas turbine engine and method according to example aspects of thepresent disclosure can increase the operability of the aircraft's gasturbine engine by making a real-time assessment of airflow distortion inthe engine airflow path of the gas turbine engine and adjusting theairflow distortion in the engine airflow path by operating a translatinginlet assembly to vary the amount and/or speed of airflow into theengine airflow path of the gas turbine engine based on the airflowdistortion assessment. Real-time pressure measurements obtained from theengine airflow path can be used to make an assessment of airflowdistortion conditions in the gas turbine engine. Airflow distortion inthe engine airflow path can then be adjusted by opening or closing thetranslating inlet assembly to increase or decrease the amount of airflowand/or airflow speed into the engine airflow path. By adjusting theairflow distortion in the engine airflow path, the operability of thegas turbine engine can be improved.

According to particular aspects of the present disclosure, a translatinginlet assembly can be used to vary the amount and/or speed of airflowentering the engine airflow path in a gas turbine engine in response toairflow distortion in the engine airflow path. The gas turbine enginecan include a compressor section, a combustion section, and a turbinesection series flow and enclosed by a casing. The gas turbine engine caninclude a translating inlet assembly at a forward end of the casing. Anengine airflow can enter the gas turbine engine through an inlet to thecompressor section between a front edge of the translating inletassembly and an inner flowpath surface. The engine airflow can then flowthrough the engine airflow path of the gas turbine engine. The frontedge of the translating inlet assembly can be configured to move betweena first position and a second position in order to vary the area of theinlet to the compressor section. For example, in an embodiment, thefront edge of the translating inlet assembly can move generally along anaxial direction in order to increase or decrease the inlet area aboutthe circumference of the inlet of the compressor section of the gasturbine engine. In another embodiment, the front edge of the translatinginlet assembly can move between a first position and a second positionsuch that the front edge of the translating inlet assembly translates atleast partially along a radial direction to increase or decrease theinlet area. In an embodiment, the translating inlet assembly can movebetween the first position and the second position in a generallyuniform manner about the circumference of the gas turbine engine suchthat the inlet area of the compressor section is generally uniform aboutthe entire circumference of the inlet. In another embodiment, thetranslating inlet assembly can move to one or more intermediatepositions such that the inlet area in the one or more intermediatepositions is less than the inlet area in the first position and morethan the inlet area in the second position.

In an embodiment, the translating inlet assembly can be configured to becontrolled to move between positions in response to airflow distortion.For example, in one embodiment, one or more pressure sensing devices canbe integrated into various components that extend into the engineairflow path of the gas turbine engine. A distortion conditionassessment can be made based on the real-time pressure measurementsobtained from the pressure sensing devices. For example, a non-uniformpressure profile across the engine airflow path can indicate thatairflow distortion is present in the engine airflow path. Thetranslating inlet assembly can then be controlled to adjust the airflowdistortion associated with the engine airflow path. For example, thetranslating inlet assembly can be controlled to increase or reduce theinlet area of the compressor section to allow or restrict airflow intothe engine airflow path, thereby reducing the airflow distortion.

In this way, the gas turbine engine and method according to exampleaspects of the present disclosure can have a technical effect ofadjusting the airflow distortion of the gas turbine engine based onreal-time airflow distortion conditions. Further, this can allow anincrease in the operability of the gas turbine engine by increasing thestall margin headroom available for operational safety.

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first”, “second”, and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows, and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexample embodiment of the present disclosure. More particularly, for theembodiment of FIG. 1, the gas turbine engine is a high-bypass turbofanjet engine 10, referred to herein as “gas turbine engine 10.” Exampleaspects of the present disclosure can be used with other suitable gasturbine engines without deviating from the scope of the presentdisclosure.

As shown in FIG. 1, the gas turbine engine 10 defines an axial directionA (extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. The gas turbine engine 10 alsodefines a circumferential direction (not depicted). In general, the gasturbine engine 10 includes a fan section 14 and a core engine 16, thefan section 14 configured in mechanical communication and positioned inflow communication with the core engine 16.

The example core engine 16 depicted generally includes a substantiallytubular outer casing 18 which defines a forward end that includes atranslating inlet assembly 112. Translating inlet assembly 112 and aninner flowpath surface 114 together define an annular inlet 20. Theouter casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

Additionally, for the embodiment depicted, the fan section 14 includes avariable pitch fan 38 having a plurality of fan blades 40 coupled to adisk 42 in a spaced apart manner. As depicted, the fan blades 40 extendoutwardly from the disk 42 generally along the radial direction R. Thefan blades 40 and disk 42 are together rotatable about the longitudinalcenterline 12 by LP shaft 36 across a power gear box 44. The power gearbox 44 includes a plurality of gears for adjusting the rotational speedof the LP shaft 36. Additionally, for the embodiment depicted, the disk42 of the variable pitch fan 38 is covered by a rotatable front hub 46aerodynamically contoured to promote an airflow through the plurality offan blades 40.

Referring still to FIG. 1, the example gas turbine engine 10additionally includes a plurality of circumferentially-spaced outletguide vanes 50. The plurality of outlet guide vanes 50 are positioneddownstream from the fan 38 along the axial direction A and extendoutwardly from the outer casing 18 of the core engine 16 generally alongthe radial direction R. Each outlet guide vane 50 defines a center ofpressure 52 (shown in FIG. 2) and a pitch axis P extending substantiallyparallel to the radial direction R. Notably, for the embodimentdepicted, the gas turbine engine 10 does not include any outer casingenclosing the fan section 14 and/or outlet guide vanes 50. Accordingly,for the embodiment depicted, the gas turbine engine 10 may be referredto as an un-ducted single fan gas turbine engine 10.

For the example gas turbine engine 10 depicted, the fan section 14, ormore particularly, the rotation of the fan blades 40 of the fan section14, provides a majority of the propulsive thrust of the gas turbineengine 10. Additionally, the plurality of outlet guide vanes 50 areprovided to increase an efficiency of the fan section 14 as well as toprovide other benefits, such as, for example, decreasing an amount ofnoise generated by the gas turbine engine 10.

During operation of the gas turbine engine 10, a volume of air 56 passesover the plurality of blades 40 of the fan section 14. A first portionof the volume of air 56, i.e., the first portion of air 60, is directedor routed through annular inlet 20 into an engine airflow path 64extending through the compressor section, the combustion section 26, theturbine section, and the exhaust section 32. The first portion of air 60may also be referred to as an engine airflow. Additionally, a secondportion of the volume of air 56, e.g., a second portion of air 62, flowsaround the core engine 16, bypassing the core engine 16. The secondportion of air 62 may also be referred to as a bypass airflow. The ratiobetween the second portion of air 62 and the first portion of air 60 iscommonly known as a bypass ratio.

Referring still to FIG. 1, the pressure of the first portion of air 60is increased as it is routed through the LP compressor 22 andsubsequently through the HP compressor 24. The compressed first portionof air 60 is then provided to the combustion section 26, where it ismixed with fuel and burned to provide combustion gases 74. Thecombustion gases 74 are routed through the HP turbine 28 where a portionof thermal and/or kinetic energy from the combustion gases 74 isextracted via sequential stages of HP turbine stator vanes 76 that arecoupled to the outer casing 18 and HP turbine rotor blades 78 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 74 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 74 via sequential stages of LP turbine stator vanes 80that are coupled to the outer casing 18 and LP turbine rotor blades 82that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38. The combustion gases 74 aresubsequently routed through the jet exhaust nozzle section 32 of thecore engine 16 to provide propulsive thrust to supplement propulsivethrust provided by the fan section 14.

Referring still to FIG. 1, downstream of an annular inlet 20 is one ormore inlet guide vanes 100. In certain example embodiments, inlet guidevane 100 may be configured to open or close, thereby restricting theflow of the first portion of air 60 into the engine airflow path 64extending through the compressor section. In certain exampleembodiments, inlet guide vane 100 can be an instrumented guide vane 400according to example embodiments of the present disclosure as depicted,for instance, in FIGS. 8 and 9.

Downstream of inlet guide vane 100 is one or more struts 102 configuredto mechanically couple outer casing 18 to the core engine 16. Strut 102extends into the engine airflow path 64 where first portion of air 60flows over strut 102. In certain example embodiments, strut 102 isconfigured to obtain pressure measurements. Downstream of strut 102 isone or more variable guide vanes 104. Variable guide vanes 104 areconfigured to open or close, thereby restricting the flow of the firstportion of air 60 into the engine airflow path 64 extending through thecompressor section. In certain example embodiments, variable guide vane104 can be an instrumented variable guide vane 400 according to exampleembodiments of the present disclosure as shown, for instance, in FIGS. 8and 9.

Referring now to FIGS. 2 and 3, a close-up, cross-sectional view of theforward end of the example gas turbine engine 10 of FIG. 1 according toexample aspects of the present disclosure is shown. As shown, the gasturbine engine 10 includes at least one control mechanism 106 configuredto adjust a variable guide vane 104. In certain example embodiments, thegas turbine engine 10 may include a plurality of control mechanisms 106,each individual control mechanism 106 configured to adjust an individualvariable guide vane 104 or other member of the airflow path.

Referring still to FIGS. 2 and 3, translating inlet assembly 112 islocated on the forward end of outer casing 18. As shown in FIGS. 2 and3, core engine 16 can include an inner flowpath surface 114. Translatinginlet assembly 112 can include a front edge 116 which can be moved byone or more actuators 118. In an embodiment, a plurality of actuators118 can be spaced about the circumferential direction of gas turbineengine 10 to move front edge 116. Together, front edge 116 and innerflowpath surface 114 can define annular inlet 20, as shown in FIG. 1.

As illustrated in FIGS. 2 and 3, in some embodiments, front edge 116 canbe moved at least partially along radial direction R between a firstposition and a second position. For example, as shown in FIG. 2, frontedge 116 is in a first position such that front edge 116 defines a firstradius 120. As shown in FIG. 3, front edge 116 is in a second positionsuch that front edge 116 defines a second radius 122. As depicted inFIGS. 2 and 3, first radius 120 is greater than second radius 122.

As depicted in FIGS. 2 and 3, first portion of air 60 enters core engine16 by passing between front edge 116 and inner flowpath surface 114. Asdepicted in FIGS. 2 and 3, front edge 116 in the first position allowsan increased first portion of air 60 to enter engine flowpath 64 ascompared to front edge 116 in the second position. In another embodimentnot shown in FIGS. 2 and 3, front edge 116 can be moved to one or moreintermediate positions such that the first portion of air 60 in theintermediate position is less than the first portion of air 60 in thefirst position and more than the first portion of air 60 in the secondposition.

In an example embodiment, front edge 116 can be moved in response to asignal from a control device, such as, for instance, control device 1000depicted in FIG. 10. For example, in an embodiment, front edge 116 canbe moved in response to a control signal sent to one or more actuators118 that controls the one or more actuators 118 to move front edge 116to a particular setpoint, such as the first, second, or intermediatepositions. In this way, the amount of airflow and/or the speed ofairflow that flows through engine airflow path 64 can be varied, therebyallowing the performance of annular inlet 20 to be varied to match theairflow demand of the LP compressor 22. In an embodiment, translatinginlet assembly 112 can be controlled to move front edge 116 to movebetween the first, second, and intermediate positions to allow theperformance of annular inlet 20 to be varied to adjust airflowdistortion to improve the operability of gas turbine engine 10, orincrease the stall margin of LP compressor 22. For example, in certainembodiments, adjusting the airflow distortion may include adjusting aninlet airflow distortion and/or adjusting an airflow mismatch and/orairflow speed mismatch between LP compressor 22 and HP compressor 24.Airflow and/or airflow speed mismatch can occur because fans, such asvariable pitch fan 38 in gas turbine engine 10, may operate within anarrow speed range, which may be elevated. In such configurations, lowpressure compressors, such as LP compressor 22, may operate at asimilarly elevated speed because they are mechanically coupled to thevariable pitch fan through a gearbox. In low power conditions, a lowpressure compressor, such as LP compressor 22, may pump too much airflowfor a high pressure compressor, such as HP compressor 24, to accept.

Referring now to FIGS. 4 and 5, a schematic, axial view of the exampleforward end of an example gas turbine engine 10 depicted in FIGS. 2 and3 is shown. As shown in FIGS. 4 and 5, gas turbine engine 10 includesouter casing 18. Outer casing 18 includes translating inlet assembly 112on a forward end. Translating inlet assembly 112 can include a frontedge 116. Gas turbine engine 10 can also include an inner flowpathsurface 114. Together, front edge 116 and inner flowpath surface 114 candefine annular inlet 20.

Front edge 116 can move between the first position, as depicted in FIG.4, and the second position, as depicted in FIG. 5. In an embodiment,front edge 116 can move between the first position and the secondposition in a substantially uniform manner about the circumferentialdirection of the gas turbine engine 10. For example, as shown in FIGS. 4and 5, front edge 116 can move between the first position and secondposition such that annular inlet 20 is generally uniform around thecircumferential direction of the gas turbine engine 10. For example,front edge 116 could comprise a plurality of overlapping panels that canopen or close to move between the first position and second position.Other embodiments of front edge 116 that are moveable between the firstposition and second position in a generally uniform manner about thecircumferential direction of gas turbine engine 10 can be used withoutdeparting from the scope or spirit of the present disclosure.Additionally, in some embodiments, front edge 116 can move to the one ormore intermediate positions in a substantially uniform manner about thecircumferential direction of the gas turbine engine 10.

As depicted in FIG. 4, front edge 116 in the first position defines thefirst radius 120 and a first circumference 124. As depicted in FIG. 5,front edge 116 in the second position defines the second radius 122 anda second circumference 126. As depicted in FIGS. 4 and 5, first radius120 is greater than second radius 122 and first circumference 124 isgreater than second circumference 126. Further, as depicted in FIG. 4,front edge 116 and inner flowpath surface 114 together define a firstarea 128 of annular inlet 20. As depicted in FIG. 5, front edge 116 andinner flowpath surface 114 together define a second area 130 of annularinlet 20. As depicted in FIGS. 4 and 5, first area 128 is greater thansecond area 130. During operation of gas turbine engine 10, front edge116 in the first position will allow an increased first portion of air60 to enter annular inlet 20 as compared to the first portion of air 60that can enter annular inlet 20 when front edge 116 is in the secondposition. Additionally and/or alternatively, front edge 116 in the firstposition will allow first portion of air 60 to enter annular inlet 20 ata first airflow speed and front edge 116 in the second position willallow the first portion of air 60 to enter annular inlet 20 at a secondairflow speed. The airflow speed of the first portion of air 60 in thesecond position is greater than the airflow speed of the first portionof air 60 in the first position. In another embodiment not shown inFIGS. 4 and 5, front edge 116 can be adjusted to one or moreintermediate positions such that the inlet area in the intermediateposition is less than the first area 128 and greater than the secondarea 130. During operation of gas turbine engine 10, front edge 116 inthe intermediate position will allow a first portion of air 60 to enterthe engine airflow path 64 that is less than the first portion of air 60in the first position and more than the first portion of air 60 in thesecond position. Additionally and/or alternatively, front edge 116 inthe intermediate position will allow a first portion of air 60 to enterannular inlet 20 at an airflow speed that is less than the airflow speedof the first portion of air 60 in the second position and greater thanthe airflow speed of the first portion of air 60 in the first position.

Referring now to FIGS. 6 and 7, the same forward end of gas turbineengine 10 depicted in FIGS. 2 and 3 is shown according to additionalexample aspects of the present disclosure. Structures that are the sameor similar to those shown in FIGS. 2 and 3 are referred to with the samereference numerals.

As depicted in FIGS. 6 and 7, translating inlet assembly 112 is locatedon the forward end of outer casing 18. As shown in FIGS. 6 and 7, coreengine 16 can include an inner flowpath surface 114. Translating inletassembly 112 can include a front edge 116 which can be moved by one ormore actuators 118. In an embodiment, a plurality of actuators 118 canbe spaced about the circumferential direction of gas turbine engine 10to move front edge 116. Together, front edge 116 and inner flowpathsurface 114 can define annular inlet 20, as shown in FIG. 1.

As illustrated in FIGS. 6 and 7, in some embodiments, front edge 116 canbe moved at least partially along axial direction A between a firstposition and a second position. For example, as shown in FIG. 6, frontedge 116 is in a first position such that annular inlet 20 has a firstinlet area 128. As shown in FIG. 7, front edge 116 is in a secondposition such that annular inlet 20 has a second inlet area 130. Asshown in FIG. 6, front edge 116 is in a first position such that frontedge 116 defines a first radius 120. As shown in FIG. 7, front edge 116is in a second position such that front edge 116 defines a second radius122. As depicted in FIGS. 6 and 7, first radius 120 and second radius122 are equal. In another embodiment, first radius 120 can be differentthan second radius 122.

As depicted in FIGS. 6 and 7, first portion of air 60 enters core engine16 by passing between front edge 116 and inner flowpath surface 114. Asdepicted in FIGS. 6 and 7, front edge 116 in the first position allowsan increased first portion of air 60 to enter engine flowpath 64 ascompared to front edge 116 in the second position. Additionally and/oralternatively, front edge 116 in the first position will allow firstportion of air 60 at a first airflow speed and front edge 116 in thesecond position will allow the first portion of air 60 at a secondairflow speed. The airflow speed of the first portion of air 60 in thesecond position is greater than the airflow speed of the first portionof air 60 in the first position. In another embodiment not shown inFIGS. 6 and 7, front edge 116 can be moved to one or more intermediatepositions such that the first portion of air 60 in the intermediateposition is less than the first portion of air 60 in the first positionand more than the first portion of air 60 in the second position.Additionally and/or alternatively, front edge 116 in the intermediateposition will allow a first portion of air 60 at an airflow speed thatis less than the airflow speed of the first portion of air 60 in thesecond position and greater than the airflow speed of the first portionof air 60 in the first position.

In an embodiment, front edge 116 can move between the first position,second position, and intermediate position in a substantially uniformmanner about the circumferential direction of the gas turbine engine 10.For example, front edge 116 can move along the axial direction betweenthe first, second, and intermediate positions such that annular inlet 20is generally uniform around the circumferential direction of the gasturbine engine 10. For example, front edge 116 could comprise a solidfront edge configured to move along the axial direction A and maintain afixed radius in the first, second, and intermediate positions. Otherembodiments of front edge 116 that can move between the first, second,and intermediate positions in a generally uniform manner about thecircumferential direction of gas turbine engine 10 can be used withoutdeparting from the scope or spirit of the present disclosure.

In an example embodiment, front edge 116 can be moved in response to asignal from a control device, such as, for instance, control device 1000depicted in FIG. 10. For example, in an embodiment, front edge 116 canbe moved in response to a control signal sent to one or more actuators118 that controls the one or more actuators 118 to move front edge 116to a particular setpoint, such as the first, second, or intermediatepositions. In this way, the amount and/or speed of airflow that flowsthrough engine airflow path 64 can be varied, and the performance ofannular inlet 20 can be varied to match the airflow demand of the LPcompressor 22. In an embodiment, translating inlet assembly 112 can becontrolled to move front edge 116 to move between the first, second, andintermediate positions to allow an increased or decreased first portionof air 60 to flow into the engine airflow path 64 or an increased ordecreased airflow speed of first portion of air 60 to adjust airflowdistortion to improve the operability of gas turbine engine 10 orincrease the stall margin of LP compressor 22. For example, in certainembodiments, adjusting the airflow distortion may include adjusting aninlet airflow distortion and/or adjusting an airflow or airflow speedmismatch between LP compressor 22 and HP compressor 24.

Referring generally to FIGS. 2-7, during operation of gas turbine engine10, front edge 116 of translating inlet assembly 112 can be configuredto be controlled to adjust airflow distortion in engine airflow path 64.For instance, a control device, such as control device 1000 shown inFIG. 10, can obtain measurements from one or more pressure sensordevices, and determine that an area of the engine airflow path isexperiencing airflow distortion, such as airflow separation. The controldevice can then control translating inlet assembly 112 to adjust theairflow distortion, by, for example, sending a control signal to one ormore actuators 118 to move front edge 116 between the first, second, andintermediate positions to increase or decrease the first portion of air60 or to increase or decrease an airflow speed of first portion of air60 flowing into the engine airflow path 64. In this way, airflowdistortion in a gas turbine engine can be adjusted, and moreover,reduced, thereby improving operability of the gas turbine engine.

In one embodiment, one or more pressure sensor devices can be located atleast partially within engine airflow path 64. In one embodiment, theone or more pressure sensor devices can be at least partially integratedinto one or more members in the engine airflow path 64, such as aninstrumented guide vane 400 as shown in FIGS. 8 and 9. In anotherembodiment, the one or more pressure sensor devices can be integratedinto inlet guide vane 100 or strut 102. Other pressure sensor devicescan also be used without departing from the scope or spirit of thepresent disclosure. Using measurements obtained by the one or morepressure sensor devices, an airflow distortion assessment can bedetermined.

Additionally, in another example embodiment, a plurality of pressuresensing devices can be spaced about the circumferential direction of gasturbine engine 10. For example, gas turbine engine 10 can include aplurality of instrumented guide vanes 400 spaced about thecircumferential direction of gas turbine engine 10 as depicted in FIG.8. Using measurements obtained by the pressure sensor devices, anairflow distortion assessment can be made. For example, a non-uniformpressure profile across the circumferential direction of the engineairflow path could indicate that airflow distortion is present in aparticular area of engine airflow path 64. Additionally, the one or morepressure sensor devices can be used at least in part for determining anairflow and/or airflow speed mismatch between two compressors, such asan LP compressor 22 and HP compressor 24. One or more actuators 118 thenbe controlled to move front edge 116 of translating inlet assembly 112to increase or decrease the first portion of air 60 flowing into engineairflow path 64, thereby adjusting, and further, reducing the airflowdistortion.

Referring now to FIG. 8, a schematic, axial view of an array ofindividual guide vanes 104 in the example gas turbine engine 10 of FIG.1 is shown. As shown, a plurality of individual guide vanes 104 areconfigured in a circumferential array located in the engine airflow path64 upstream of the LP Compressor 22. As depicted in FIG. 8, fiveinstrumented guide vanes 400, as discussed below in greater detail withrespect to FIG. 9, are included in the array of individual guide vanes104. As will be discussed in greater detail below with reference to FIG.9, each individual instrumented guide vane 400 is configured with apressure sensing device. As shown in FIG. 8, the pressure sensing deviceincludes one or more taps 202 extending through the individualinstrumented guide vane 400 and one or more local transducers 204configured to measure an air pressure from the one or more taps 202.However, it will be apparent to those skilled in the art will that thepressure sensing device can be any suitable device configured to sensepressure without departing from the scope or spirit of the presentdisclosure. According to example aspects of the present disclosure, oneor more pressure sensor devices, such as an array of instrumented guidevanes depicted in FIG. 8, can be used for obtaining one or moremeasurements to determine an airflow distortion condition in the engineairflow path, such as an inlet airflow distortion. Additional pressuresensor devices not depicted can also be used for obtaining measurementsto determine other airflow distortion conditions, such as an airflowand/or airflow speed mismatch between two compressors, such as a LPcompressor 22 and HP compressor 24 (see e.g., FIG. 1). As shown in FIG.8, local transducer 204 is configured to send data indicative of an airpressure to a digital communication bus 206. Digital communication bus206 then sends the data indicative of an air pressure to controller 208.In an embodiment, controller 208 can be a control device programmed toperform operations, such as control device 1000 depicted in FIG. 10.Controller 208 can control various actuators based on the dataindicative of an air pressure, such as one or more actuators 118 of atranslating inlet assembly 112.

FIG. 9 is a schematic of an individual instrumented guide vane 400 foran example gas turbine engine according to example embodiments of thepresent disclosure. Instrumented guide vane 400 can be a variable guidevane 104 or a stationary guide vane 100. As depicted in FIG. 9,instrumented guide vane 400 can be configured in a nonsymmetricalairfoil shape generally having a “tear drop” shape with a leading edge410, a pressure side 420, and a suction side 430. However, in otherexample embodiments, the instrumented guide vane 400 may instead defineany other suitable symmetrical or nonsymmetrical shape or configuration.In some implementations, leading edge 410 can be configured withinengine airflow path 64 such that first portion of air 60 flowingdownstream of annular inlet 20 first comes into contact with leadingedge 410 before flowing over pressure side 420 and suction side 430 andcontinuing into LP compressor 22.

Referring still to FIG. 9, one or more leading edge taps 412, pressureside taps 422 and/or suction side taps 432 are integrated intoinstrumented guide vane 400. The leading edge taps 412, pressure sidetaps 422, and suction side taps 432 are depicted in phantom. As depictedin FIG. 9, two leading edge inlets 414 are spaced radially along leadingedge 410 to allow air from an engine airflow, such as first portion ofair 60, to enter leading edge inlet 414 and flow through leading edgetap 412 to a local transducer 204 (not shown in FIG. 9). In anotherembodiment, a single leading edge inlet 414 and leading edge tap 412 canbe integrated into leading edge 410. In another embodiment three or moreleading edge inlets 414 and leading edge taps 412 can be integrated intoleading edge 410.

Referring still to FIG. 9, two pressure side inlets 424 are spacedaxially along pressure side 420 to allow an engine airflow, such as airfrom first portion of air 60, to enter pressure side inlet 424 and flowthrough pressure side tap 422 to a local transducer 204 (not shown inFIG. 9). In another embodiment, a single pressure side inlet 424 andpressure side tap 422 are integrated into pressure side 420. In anotherembodiment three or more pressure side inlets 424 and pressure side taps422 are integrated into pressure side 420.

Referring still to FIG. 9, two suction side inlets 434 are spacedaxially along suction side 430 to allow air from an engine airflow, suchas first portion of air 60, to enter suction side inlet 434 and flowthrough suction side tap 432 to a local transducer 204 (not shown inFIG. 9). The suction side inlets 434 are depicted in phantom. In anotherembodiment a single suction side inlet 434 and suction side tap 432 areintegrated into suction side 430. In another embodiment, three or moresuction side inlets 434 and suction side taps 432 are integrated intosuction side 430.

Referring still to FIG. 9, in an embodiment, local transducer 204 (notshown) can be configured to provide measurements of a pressuredifferential between a pressure side tap 422 and a suction side tap 432.In another embodiment, local transducer 204 (not shown) can beconfigured to provide measurements of absolute pressures from a pressureside tap 422 and a suction side tap 432.

FIG. 10 depicts an example control device used in a control systemaccording to example embodiments of the present disclosure. As shown,the control device(s) 1000 can include one or more processors 1012 andone or more memory devices 1014. The one or more processors 1012 caninclude any suitable processing device, such as a microprocessor,microcontroller, integrated circuit, logic device, or other suitableprocessing device. The one or more memory devices 1014 can include oneor more computer-readable media, including, but not limited to,non-transitory computer-readable media, RAM, ROM, hard drives, flashdrives, or other memory devices.

The one or more memory devices 1014 can store information accessible bythe one or more processors 1012, including computer-readableinstructions 1016 that can be executed by the one or more processors1012. The instructions 1016 can be any set of instructions that whenexecuted by the one or more processors 1012, cause the one or moreprocessors 1012 to perform operations. The instructions 1016 can beimplemented in software written in any suitable programming language orcan be implemented in hardware. In some embodiments, the instructions1016 can be executed by the one or more processors 1012 to cause the oneor more processors 1012 to perform operations, such as the operationsfor controlling a translating inlet assembly to adjust airflowdistortion in a gas turbine engine as described with reference to FIG.11.

Referring to FIG. 10, the memory devices 1014 can further store data1018 that can be accessed by the processors 1012. The data 1018 caninclude, for instance, operating parameters, pressure measurementsobtained from the engine airflow path, and other data. The data 1018 canalso include data associated with models and algorithms used to performthe example methods according to example aspects of the presentdisclosure, such as models and algorithms for determining a distortioncondition.

The control device(s) 1000 can further include a communicationsinterface 1020. The communications interface 1020 can be configured tocommunicate with aircraft systems over a communication network 1040. Forinstance, the communications interface 1020 can receive data indicativeof a pressure obtained by a pressure sensing device, such as a tap 202and local transducer 204. In one embodiment, the communicationsinterface 1020 can provide control commands to an engine control system1050 that has one or more actuators to control various components of thegas turbine engine 10, such as, but not limited to, an actuator 118 of atranslating inlet assembly 112. The communications interface 1020 caninclude any suitable components for interfacing with one more otherdevices, including for example, transmitters, receivers, ports,controllers, antennas, or other suitable components.

The technology discussed herein makes computer-based systems, as well asactions taken and information sent to and from such systems. One ofordinary skill in the art will recognize that the inherent flexibilityof computer-based systems allows for a great variety of possibleconfigurations, combinations, and divisions of tasks and functionalitybetween and among components. For instance, processes discussed hereinmay be implemented using a single computing device or multiple computingdevices working in combination. Databases, memory, instructions, andapplications may be implemented on a single system or distributed acrossmultiple systems. Distributed components may operate sequentially or inparallel.

Referring now to FIG. 11, a flow diagram of an example method (1100)according to example embodiments of the present disclosure is depicted.FIG. 11 can be implemented by one or more control devices, such as thecontrol device 1000 depicted in FIG. 10. In addition, FIG. 11 depictssteps performed in a particular order for purposes of illustration anddiscussion. Those of ordinary skill in the art, using the disclosuresprovided herein, will understand that the various steps of any of themethods disclosed herein can be modified, adapted, expanded, rearrangedand/or omitted in various ways without deviating from the scope of thepresent disclosure.

At (1102), the method can include obtaining one or more measurementsfrom one or more pressure sensor devices. The one or more measurementscan be obtained by, for example, a local transducer 204 operativelyconnected to an instrumented guide vane 400 as shown in FIGS. 8 and 9.Alternatively, the one or more measurements can be obtained from anyother suitable pressure sensor device.

At (1104), the method can include determining a distortion conditionassociated with the engine airflow path of a gas turbine engine from theone or more measurements. For example, using the one or moremeasurements, a distortion condition can be determined, such as a localflow disruption in the engine airflow path 64 of the gas turbine engine10 that causes portions of the LP compressor 22 to be at or belowpressures sufficient to cause stall conditions.

At (1106), the method can include determining a control signal foractivation of a translating inlet assembly based at least in part on thedistortion condition assessment. For example, a set point can bedetermined for a translating inlet assembly 112 to increase the airflowthrough the translating inlet assembly 112 in order to energize an areaof the engine airflow path 64 that is experiencing a local flowdisruption. A control signal representing the determined set point ofthe translating inlet assembly 112 can then be sent to one or moreactuators 118 in order to adjust the translating inlet assembly 112.

At (1108), the method can include controlling the translating inletassembly 112 to adjust the distortion condition based on the controlsignal. For example, one or more actuators 122 can move front edge 116of translating inlet assembly 112 based on the control signal. Thetranslating inlet assembly 112 can be then controlled to move to thedetermined set point to adjust the airflow distortion. In this way, atranslating inlet assembly 112 can adjust the airflow distortionassociated with the gas turbine engine.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A core engine for a gas turbine engine,comprising: a compressor section, a combustion section, and a turbinesection defining at least in part an engine airflow path for the coreengine; an inner flowpath surface positioned at least partially withinthe compressor section and defining at least in part the engine airflowpath; a core casing at least partially enclosing the compressor sectionand defining a forward end, the core casing comprising a translatinginlet assembly at the forward end, the translating inlet assembly andinner flowpath surface together defining an inlet to the compressorsection, the translating inlet assembly moveable between a firstposition defining a first inlet area at the inlet and a second positiondefining a second inlet area at the inlet, the first inlet area beinggreater than the second inlet area.
 2. The core engine of claim 1,wherein the translating inlet assembly allows a first engine airflow ata first airflow speed into the engine airflow path in the firstposition, wherein the translating inlet assembly allows a second engineairflow at a second airflow speed into the engine airflow path in thesecond position, wherein at least one of the first engine airflow isgreater than the second engine airflow or the first engine airflow speedis less than the second engine airflow speed.
 3. The core engine ofclaim 1, wherein the translating inlet assembly is configured to becontrolled based on an airflow distortion in the engine airflow path,and wherein the airflow distortion is an inlet airflow distortion. 4.The core engine of claim 1, wherein the compressor section includes afirst compressor and a second compressor, wherein the translating inletassembly is configured to be controlled based on an airflow distortionin the engine airflow path, and wherein the airflow distortion is anairflow mismatch between the first and second compressors.
 5. The coreengine of claim 1, wherein the translating inlet assembly is configuredto be controlled based on an airflow distortion in the engine airflowpath, the core engine further comprising: one or more pressure sensordevices located at least partially in the engine airflow path forobtaining one or more measurements associated with airflow distortion;wherein the translating inlet assembly is configured to be controlledbased at least in part on signals from the one or more pressure sensordevices.
 6. The core engine of claim 1, wherein the core engine definesa circumferential direction, wherein the translating inlet assembly ismovable between the first position and the second position in asubstantially uniform manner about the circumferential direction of thecore engine.
 7. The core engine of claim 1, wherein the translatinginlet assembly is additionally movable to an intermediate position,wherein the intermediate position defines an intermediate inlet area,wherein the intermediate inlet area is less than the first inlet areaand greater than the second inlet area.
 8. The core engine of claim 7,wherein the translating inlet assembly is movable between the first,second, and intermediate positions based on signals from a controller toadjust an airflow distortion in the engine airflow path.
 9. The coreengine of claim 1, wherein the translating inlet assembly defines afront edge, wherein the core engine defines an axial direction, whereinthe front edge of the translating inlet assembly moves between the firstposition and the second position at least partially along the axialdirection.
 10. The core engine of claim 1, wherein the translating inletassembly defines a front edge, wherein the core engine defines acircumferential direction, wherein the front edge defines a firstcircumference in the first position, wherein the front edge defines asecond circumference in the second position, wherein the secondcircumference is less than the first circumference.
 11. A method foradjusting airflow distortion in a gas turbine engine on an aircraft, thegas turbine engine comprising a compressor section, a combustionsection, and a turbine section in series flow, the compressor section,combustion section, and turbine section defining at least in part anengine airflow path, the gas turbine engine further comprising an innerflow path surface positioned at least partially within the compressorsection and defining at least in part the engine airflow path, the gasturbine engine further comprising a core casing at least partiallyenclosing the compressor section and defining a forward end, the methodcomprising: determining, by one or more control devices, an airflowdistortion condition associated with the engine airflow path; andcontrolling, by the one or more control devices, a translating inletassembly to adjust the airflow distortion condition of the gas turbineengine, wherein the core casing comprises the translating inlet assemblyat the forward end, the translating inlet assembly and inner flow pathsurface together defining an inlet to the compressor section, thetranslating inlet assembly moveable between a first position defining afirst inlet area and a second position defining a second inlet area, thefirst inlet area being greater than the second inlet area.
 12. Themethod of claim 11, wherein the translating inlet assembly isadditionally movable to an intermediate position, wherein theintermediate position defines an intermediate inlet area, wherein theintermediate inlet area is less than the first inlet area and greaterthan the second inlet area.
 13. The method of claim 12, wherein thetranslating inlet assembly allows a first engine airflow into the engineairflow path in the first position, wherein the translating inletassembly allows a second engine airflow into the engine airflow path inthe second position, wherein the translating inlet assembly allows anintermediate engine airflow into the engine airflow path in theintermediate position, wherein the intermediate engine airflow is lessthan the first engine airflow and greater than the second engineairflow.
 14. The method of claim 12, wherein controlling the translatinginlet assembly comprises moving the translating inlet assembly betweenthe first, second, and intermediate positions.
 15. The method of claim11, wherein controlling the translating inlet assembly to adjust theairflow distortion condition comprises controlling the translating inletassembly to reduce the airflow distortion condition.
 16. The method ofclaim 11, wherein determining the airflow distortion conditionassociated with the engine airflow path comprises obtaining one or moremeasurements using one or more pressure sensor devices, whereincontrolling the translating inlet assembly to adjust the airflowdistortion comprises controlling the translating inlet assembly based atleast in part on the one or more measurements obtained using the one ormore pressure sensor devices.
 17. The method of claim 11, wherein thetranslating inlet assembly defines a front edge, wherein the core enginedefines an axial direction, wherein the front edge of the translatinginlet assembly moves between the first position and the second positionalong the axial direction.
 18. The method of claim 11, wherein thetranslating inlet assembly defines a front edge, wherein the core enginedefines a circumferential direction, wherein the front edge defines afirst circumference in the first position, wherein the front edgedefines a second circumference in the second position, wherein thesecond circumference is less than the first circumference.
 19. A gasturbine engine system for an aircraft comprising: a compressor section,a combustion section, and a turbine section defining at least in part anengine airflow path for the core engine; an inner flowpath surfacepositioned at least partially within the compressor section and definingat least in part the engine airflow path; a core casing at leastpartially enclosing the compressor section and defining a forward end,the core casing comprising a translating inlet assembly at the forwardend, the translating inlet assembly and inner flowpath surface togetherdefining an inlet to the compressor section, the translating inletassembly moveable between a first position defining a first inlet areaat the inlet and a second position defining a second inlet area at theinlet, the first inlet area being greater than the second inlet area;and a controller operably connected to the translating inlet assembly,the controller comprising one or more processors and one or more memorydevices located on an aircraft, the one or more memory devices storinginstructions that when executed by the one or more processors cause theone or more processors to perform operations, the operations comprising:determine an airflow distortion condition within the engine airflowpath; and control the translating inlet assembly to adjust an airflowthrough the airflow passage to adjust the determined airflow distortioncondition.
 20. The gas turbine engine system of claim 19, wherein thecore engine comprises one or more pressure sensor devices located atleast partially in the engine airflow path configured to obtainmeasurements to determine the airflow distortion condition, wherein thetranslating inlet assembly is controlled by the controller based atleast in part on measurements obtained by the one or more pressuresensors.